Aircraft gas turbine engines are required to operate over a wide variety of inlet fuel conditions depending upon a number of parameters that affect the inlet. Inlet fuel conditions are most noticeably affected by loss of airframe boost pressure, or by a desire to operate in a suction-feed fuel delivery mode. Inlet fuel is also affected by fuel temperature, fuel true vapor pressure, line loss and fuel flow rate. Two primary design parameters are vapor/liquid ratio at the pump inlet and net positive suction pressure or NPSP, which is the pressure at the pump inlet above true vapor pressure of the fuel at the inlet temperature. System design specifications typically require fuel pumps to operate at a specified flow rate with a vapor/liquid inlet ratio of 0.45, and with an NPSP of 5.0 psi. Newer system specifications, however, typically require the 0.45 vapor/liquid inlet ratio capability over a wider fuel flow range, and may even require a 1.0 vapor/liquid ratio with intermittent all-liquid or all-vapor operation. Furthermore, the NPSP requirements have been increased typically to 5.0 psi over the entire fuel flow range, and in some cases even 1.0 psi over the entire flow range.
It is therefore a general object of the present invention to provide a hydraulic fluid pump that is capable of satisfying flow requirements in aircraft turbine engine fuel delivery systems over an extended engine operating range, and that is adapted to operate at a vapor/liquid inlet ratio of up to 1.0 without cavitation and at 2.0 psi NPSP over an extended fuel flow range. A further object of the present invention is to provide a fuel pump of the described character that is economical and efficient in construction in terms of the stringent weight and volume requirements of aircraft applications, and that provides reliable service over an extended operating life.